EP3059393A1 - Turbosoufflante à engrenage à rapport de transmission élevé et une capacité à haute température - Google Patents
Turbosoufflante à engrenage à rapport de transmission élevé et une capacité à haute température Download PDFInfo
- Publication number
- EP3059393A1 EP3059393A1 EP16156289.7A EP16156289A EP3059393A1 EP 3059393 A1 EP3059393 A1 EP 3059393A1 EP 16156289 A EP16156289 A EP 16156289A EP 3059393 A1 EP3059393 A1 EP 3059393A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- blades
- rotor
- fan
- turbine rotor
- gas turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
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- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 29
- 230000009467 reduction Effects 0.000 claims abstract description 14
- 230000002708 enhancing effect Effects 0.000 claims abstract description 5
- 238000001816 cooling Methods 0.000 claims description 7
- 239000013078 crystal Substances 0.000 claims description 6
- 239000000463 material Substances 0.000 claims description 6
- 241000237503 Pectinidae Species 0.000 description 5
- 239000000446 fuel Substances 0.000 description 5
- 235000020637 scallop Nutrition 0.000 description 5
- 230000001965 increasing effect Effects 0.000 description 4
- 230000008901 benefit Effects 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 3
- 230000003068 static effect Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
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- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004044 response Effects 0.000 description 1
Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/16—Control of working fluid flow
- F02C9/18—Control of working fluid flow by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/286—Particular treatment of blades, e.g. to increase durability or resistance against corrosion or erosion
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/10—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/32—Arrangement, mounting, or driving, of auxiliaries
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/10—Metals, alloys or intermetallic compounds
- F05D2300/17—Alloys
- F05D2300/175—Superalloys
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/607—Monocrystallinity
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This application relates to a geared turbofan engine which may be particularly beneficial for application on regional jet engines.
- Gas turbine engines typically include a fan delivering air into a compressor, and into a bypass duct as propulsion air. Air in the compressor is compressed and delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors driving them to rotate.
- a turbine rotor drove an upstream compressor rotor and a fan rotor at a single speed.
- a gas turbine engine comprises a fan rotor having blades with an outer diameter.
- the outer diameter is greater than or equal to 77 inches (196 centimeters) and less than or equal to 135 inches (343 centimeter).
- the fan rotor has less than or equal to 26 fan blades, and is driven by a fan drive turbine through a gear reduction.
- the gear reduction has a gear ratio of greater than 2.6:1.
- the fan rotor delivers air into a bypass duct as bypass air, and into a duct leading to a compressor rotor as core air.
- a ratio of bypass air to the core air is greater than or equal to 12:1.
- An upstream turbine rotor is upstream of the fan drive turbine and drives a compressor rotor.
- the upstream turbine rotor has at least two stages, and the fan drive turbine rotor has at least three stages.
- the turbine blades in at least one stage of the fan drive turbine rotor are provided with a performance enhancing feature.
- the performance enhancing feature is at least one of the blades being manufactured by a directionally solidified blade material.
- the blades are provided as single crystal blades, and have a radially outer platform provided with scalloping to reduce the weight of the blades.
- the blades are provided with cooling air.
- the gear ratio is greater than 3.06:1.
- the gear ratio is a star gearbox.
- an area is defined at a downstream end of the upstream turbine rotor, and a second area is defined at a downstream end of the fan drive turbine rotor.
- the upstream turbine rotor rotates ng at a first speed and the fan drive turbine rotor rotates at a second speed.
- a performance quantity of the upstream turbine rotor is defined by the first cross-sectional area multiplied by the first speed squared.
- a performance quantity of the fan drive turbine rotor is defined by the second cross-sectional area multiplied by the second speed squared. The first performance ratio is less than the second performance ratio.
- the second performance quantity is greater than or equal to at least 5.0 times 10 squared (5.0 x 10 2 ).
- the blades in the at least one stage are manufactured by a directionally solidified blade material.
- the blades in the at least one stage are provided as single crystal blades.
- the blades in the at least one stage have a radially outer platform, which is provided with scalloping to reduce the weight of the blades in the at least one stage.
- the blades in the at least one stage are provided with cooling air.
- an area is defined at a downstream end of the upstream turbine rotor, and a second area is defined at a downstream end of the fan drive turbine rotor.
- the upstream turbine rotor rotates at a first speed and the fan drive turbine rotor rotates at a second speed.
- a performance quantity of the upstream turbine rotor is defined by the first cross-sectional area multiplied by the first speed squared and a performance quantity of the fan drive turbine rotor being defined by the second cross-sectional area multiplied by the second speed squared.
- the first performance ratio is less than the second performance ratio.
- the second performance quantity is greater than or equal to at least 5.0 times 10 squared (5.0 x 10 2 ).
- the blades in the at least one stage are manufactured by a directionally solidified blade material.
- the blades are provided as single crystal blades.
- the blades in the at least one stage have a radially outer platform, which is provided with scalloping to reduce the weight of the blades in the at least one stage.
- the blades in the at least one blade row are provided with cooling air.
- the gas turbine engine is for use on a single aisle aircraft.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- Alternative engines might include an augmentor section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
- the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46.
- the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
- the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54.
- a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54.
- a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
- the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28.
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
- the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
- the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
- gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
- the engine 20 in one example is a high-bypass geared aircraft engine.
- the engine 20 bypass ratio is greater than about six (6:1), with an example embodiment being greater than about ten (10:1)
- the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 (2.3:1)
- the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1).
- the engine 20 bypass ratio is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1).
- Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
- the fan section 22 of the engine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters).
- the flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
- "Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
- Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)] 0.5 .
- the "Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second (350.5 meters/second).
- FIG. 2 shows an alternative engine 80.
- Alternative engine 80 has a fan rotor 82 being driven by a gear reduction 83, which is, in turn, driven by a fan drive turbine 84.
- An upstream compressor rotor 86 is driven by an intermediate turbine rotor 88.
- a downstream compressor rotor 90 is driven by a high pressure turbine rotor 92.
- a combustor 94 is also shown schematically.
- the upstream turbine section as described, would be the high pressure turbine rotor 54, while the downstream turbine rotor or fan drive turbine rotor would be the low pressure turbine 46.
- the upstream turbine rotor may be the intermediate turbine rotor 88, while the downstream turbine rotor or fan drive turbine rotor would be turbine rotor 84.
- the disclosed gas turbine engine will have a very high bypass ratio.
- a gear ratio of the gear reduction will be relatively high.
- the speed of the fan drive turbine will be beneficially increased. This can be best achieved by increasing the gear reduction such that the fan rotor does not also increase in speed.
- the following disclosure will provide improvements to the fan drive turbine and the overall turbine section, to increase the ability to withstand high temperatures.
- bypass ratio is desirably greater than or equal to about 12:1.
- the gear reduction preferably is a gear ratio of 3.06:1, and is a star gearbox.
- a fan diameter is preferably greater than or equal to 77 inches (196 centimeters), but less than or equal to 135 inches (343 centimeters). Preferably, there are less than 26 fan blades.
- turbine rotor 96 which is directly upstream of the fan drive turbine rotor 98, there are preferably at least two stages.
- the fan drive turbine preferably has at least three stages.
- a first area A H is defined at a downstream end of the upstream turbine rotor 96.
- a second area A L is defined at a downstream end of the fan drive turbine rotor 98.
- the upstream turbine rotor 96 rotates at a first speed and the fan drive turbine rotor 98 rotates at a second speed.
- a first performance quantity (AN 2 ) of the upstream turbine rotor 96 is defined by cross-sectional area A H multiplied by its speed squared.
- a second performance quantity (AN 2 ) of the fan drive turbine rotor 98 is defined by cross-sectional area A L multiplied by its speed squared.
- the second performance ratio is greater than or equal to at least 5.0 times 10 squared (5.0 x 10 2 ) RPM in 2 (3226 RPM cm 2 ) at engine redline speed, which is slightly higher than takeoff.
- the performance quantity of upstream turbine rotor 58 is less than the performance quantity of the fan drive turbine rotor 98.
- an engine 110 has combustor 114.
- a first turbine vane 114 will "see” products of combustion at temperature of at least 2790°F (1532°C). This is known as the T4.1 location.
- An upstream turbine rotor has two stages, 116A and 116B.
- the fan drive turbine rotor has three stages, 118A, 118B, and 118C.
- blades 120A, 120B, and 120C are associated with the stages 118A, 118B, and 118C, respectively.
- the blades 120A, 120B, and 120C preferably have some feature provided to increase their resistance to high temperatures. Such features are disclosed in the following paragraphs and Figures 5 and 6 .
- the blades may be formed of a directionally solidified blade material, or a single crystal blade.
- a blade 121 may be utilized that has an outer platform 122, an airfoil 124, and an inner platform 126.
- Outer platform 122 is provided with scalloping to reduce the weight.
- the turbofan has a high temperature level, and everything in it is similarly hot. It is desirable to produce an engine with fan-drive turbine blades with high (for example 15000) cycles of life.
- the scallops increase to temperature tolerance.
- allowable stress which is the stress one can put onto the part at the temperature it is running.
- a limiting area of this blade is in fact at the neck, where it meets the disk, but one can reduce the stress, and raise the temperature tolerance to meet the cycle goal by putting the scallops in so the pull is less.
- leading edge 128 and a trailing edge 130 there is a leading edge 128 and a trailing edge 130.
- the trailing edge 130 may be provided with scalloping 132 on an underside.
- the platform 122 may be provided with scalloping 134 at the leading edge 128. Further scalloping 136 may be provided on an underside of the platform.
- the "scallops" could be described as an interrupted body of revolution.
- a shroud has two ID corners and two OD corners, and these are interrupted by the scallops and are no longer a two dimensional curve, but rather become a 3 dimensional curve.
- the OD surface of the shroud, adjacent to the knife edges is conventionally a surface of simple revolution also; adding scallops interrupts that surface and the surface moves toward the centerline locally.
- Blade 140 is illustrated in Figure 6 .
- Blade 140 is shown to have an airfoil 142 provided with cooling air 144.
- the provision of cooling air to a turbine blade is not, in itself, novel, however, the use of such a blade in the particular disclosed system provides benefits.
- FIG. 7 schematically shows an aircraft 148, which may particularly benefit from the disclosed engine.
- the aircraft 148 has a passenger section 150, a single aisle 152, and another passenger section 154.
- So-called "single aisle" aircraft are utilized on medium to short range flights and in the range of 3,900 miles (6.3 km) or less.
- the passenger capacity for such an aircraft is less than 300 and more narrowly less than 250.
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Materials Engineering (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP22164678.9A EP4036373A1 (fr) | 2015-02-18 | 2016-02-18 | Turboréacteur à engrenage à rapport de transmission élevé et une capacité à haute température |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US14/624,668 US20160237914A1 (en) | 2015-02-18 | 2015-02-18 | Geared Turbofan With High Gear Ratio And High Temperature Capability |
Related Child Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP22164678.9A Division EP4036373A1 (fr) | 2015-02-18 | 2016-02-18 | Turboréacteur à engrenage à rapport de transmission élevé et une capacité à haute température |
Publications (2)
Publication Number | Publication Date |
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EP3059393A1 true EP3059393A1 (fr) | 2016-08-24 |
EP3059393B1 EP3059393B1 (fr) | 2022-03-30 |
Family
ID=55453018
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP22164678.9A Withdrawn EP4036373A1 (fr) | 2015-02-18 | 2016-02-18 | Turboréacteur à engrenage à rapport de transmission élevé et une capacité à haute température |
EP16156289.7A Active EP3059393B1 (fr) | 2015-02-18 | 2016-02-18 | Turbosoufflante à engrenage à rapport de transmission élevé et une capacité à haute température |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
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EP22164678.9A Withdrawn EP4036373A1 (fr) | 2015-02-18 | 2016-02-18 | Turboréacteur à engrenage à rapport de transmission élevé et une capacité à haute température |
Country Status (2)
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US (2) | US20160237914A1 (fr) |
EP (2) | EP4036373A1 (fr) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN111350612A (zh) * | 2018-12-21 | 2020-06-30 | 劳斯莱斯有限公司 | 涡轮引擎 |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP3933181A1 (fr) * | 2013-08-20 | 2022-01-05 | Raytheon Technologies Corporation | Moteur à turbine à gaz à engrenages de poussée élevée |
GB201918778D0 (en) * | 2019-12-19 | 2020-02-05 | Rolls Royce Plc | Gas turbine engine with shaft bearings |
JP7284737B2 (ja) * | 2020-08-06 | 2023-05-31 | 三菱重工業株式会社 | ガスタービン静翼 |
US20220397062A1 (en) * | 2021-06-11 | 2022-12-15 | Raytheon Technologies Corporation | Gas turbine engine with electrically driven compressor |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
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US20020159889A1 (en) | 2001-04-27 | 2002-10-31 | Bunker Ronald Scott | Turbine blade having a cooled tip shroud |
US20090053047A1 (en) * | 2007-08-22 | 2009-02-26 | General Electric Company | Turbine bucket tip shroud edge profile |
US20130192200A1 (en) * | 2012-01-31 | 2013-08-01 | United Technologies Corporation | Geared turbofan gas turbine engine architecture |
Family Cites Families (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2293631B (en) * | 1994-09-30 | 1998-09-09 | Gen Electric | Composite fan blade trailing edge reinforcement |
US6902376B2 (en) * | 2002-12-26 | 2005-06-07 | General Electric Company | Compressor blade with dovetail slotted to reduce stress on the airfoil leading edge |
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2016
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- 2016-02-18 EP EP16156289.7A patent/EP3059393B1/fr active Active
-
2019
- 2019-04-26 US US16/395,623 patent/US20200025106A1/en not_active Abandoned
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Also Published As
Publication number | Publication date |
---|---|
US20160237914A1 (en) | 2016-08-18 |
EP4036373A1 (fr) | 2022-08-03 |
US20200025106A1 (en) | 2020-01-23 |
EP3059393B1 (fr) | 2022-03-30 |
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